Gas turbines are used in a variety of applications including aircraft power generation. At the core of such a turbine are a number of stages including a compressor that is used to increase the pressure of the incoming free stream flow. The compressor typically includes a rotor that includes a rotating hub with a number of radially extending blades. The rotor is typically found within a housing or shroud referred to as a casing, wherein the blade tips extend as close as possible to the casing “endwall”. These turbines have evolved to provide a reliable power source for aircraft, but also carry inherent limitations. One pertinent limitation is the phenomenon known as stall.
As is well known by gas turbine practitioners, stall or surge is a phenomenon that is characteristic of all types of axial or centrifugal compressors that limits their pressure rise capability. Those involved in compressor technology pay great heed to the surge characteristics of the compressors to assure proper compromise between performance and safe operation. During compressor operation, stall occurs when the stream wise momentum imparted to the air by the blades is insufficient to overcome the pressure rise across the compressor stage resulting in a reduction in airflow through a portion of the compressor stage. The flow leakage that occurs across the clearance gap between the compressor rotor blade tip and stationary casing endwall is one well known mechanism for reducing the total stream wise momentum through the blade passage, thus reducing the blade pressure rise capability and moving the compressor closer towards the stall condition. If no corrective action is taken, the compressor stall may propagate through several compressor stages, starving the gas turbine of sufficient air to maintain engine speed that decreases the turbines ability to create power, further reducing the output of the engine. Further, the instability created by stall may generate forces that can potentially damage the engine. If stall spreads to encompass all stages within the compressor, the global flow through the engine may actually be reversed resulting in the phenomena known as surge that exacerbates the losses, reduces engine power and increases the potential for catastrophic damage. To avoid stall, operating limits may be placed on the engine to define a safe operating range, where stall is unlikely. This operating range between the safe operating limit and stall is often referred to as the “stall margin.” As in many systems, greater efficiency is achieved at higher operating conditions, and, thus, to that extent, engine efficiency is sacrificed to obtain safe operating conditions. As will be appreciated, to further avoid stall and to improve engine performance, it is desirable to expand the stall margin for a given engine. The current trend towards increased pressure rise per stage and increased blade aerodynamic loading, however, tends to reduce the stable operating range of turbine compressors. To maintain adequate stall margin, the compressor must either operate in an inefficient manner i.e. further from the optimum efficiency point, or methods must be devised to extend the stable operating range of the compressor. Over the last thirty years various forms of endwall treatments have been employed for enhancing compressors stall range, generally at the expense of compressor efficiency.
The current state of the art in endwall treatment and designs utilizes the static pressure rise created at the compressor to recirculate high-pressure fluid to energize low momentum fluid along the casing or hub endwall, hereinafter referred to as endwall blockage. To energize the low momentum fluid, high-pressure fluid is channeled from the rear to the front of a compressor rotor through a path contained within the casing surrounding the compressor. The high-pressure fluid is then reinjected upstream of the rotor to energize the low momentum fluid at the casing or hub.
For example, one endwall treatment known in the industry incorporates a passage having an outlet port disposed over the tip of the blade and near the leading edge of the blade. The outlet port is disposed at an acute angle relative to the plane of the blade tip. An inlet port is located downstream of the outlet port near the trailing edge of the blade. In this design, the inlet port is located over the tip of the blade and connected to the outlet port by a passage that extends initially radially outward at an acute angle relative to the casing and then curves to form an elbow at its radial extremity and continues in angular fashion radially inward toward the outlet port. To counteract the high swirl component of air taken from the trailing edge of the blade tip, an anti-swirl element is located within the casing to de-swirl the air ingested at the inlet. The anti-swirl elements include reverse swirl vanes disposed at an angle relative to the main airflow and adapted to reorient the ingested air in a flow path parallel to the main flow. In this design it was observed that such a treatment could recover the energy of the low momentum flow leaving the rotor tip and return it to the main flow in an essentially axial direction. To achieve this, the dimension of the inlet, outlet, and passageway were designed to recirculate 12% of the total airflow in the main flow.
In another design in the industry, a similar passageway is used to remove low momentum fluid from the main flow of an aircraft engine. In this design, like the previously mentioned example, the flow is removed downstream of the leading edge of the blade's tip and returned at a point over the blade tip. In contrast to the previously discussed design, the inlet and outlet port angles extend at an oblique angle to the plane of the blade tip. A critical feature of this design is that the upper limit of the air removed is 8 percent. In a later patent, U.S. Pat. No. 5,431,533, after realizing that the recirculation of low momentum fluid still did not provide desired maintenance of engine efficiency, operation of the recirculating passage discussed in the previous example was limited to periods when incidence of stall was more likely. At all other times, the recirculation passages were blocked off by inflatable membranes located near the inlet and outlet sides of the passage.
Recognizing the difficulty of individually machining vanes capable of recirculating low momentum fluid, as described, within the casing, a more recent design known within the industry provides an annular plenum formed by the attachment of an insert to the casing's inner wall. The insert is provided with a recessed portion that is located on the radial outward surface of the insert that cooperates with the inner surface of the casing to define an annular plenum. Inlet and outlet ports extend through the insert to communicate with the plenum. These ports, as with previously described ports, extend at an oblique angle relative to the tip of the blade and are located above the blade tip.
This advancement of using a recirculated endwall treatment has provided the greatest stall range capability with the least decrement to compressor efficiency of previous endwall treatment concepts, but such treatment still results in an appreciable decrement in compressor efficiency.